1. Field of the Invention
The present invention relates generally to a turbine blade in a gas turbine engine, and more specifically to cooling of the fillet along the trailing edge of the turbine blade.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
In an industrial gas turbine engine, the hot gas flow developed within the combustor is passed through a multiple staged turbine to convert the hot gas flow into mechanical energy by rotating the shaft of the engine. It is well known, in the art of gas turbine engines, that the engine efficiency can be increased by providing for a higher gas flow temperature entering the turbine. However, the highest temperature that can be passed into the turbine is depended upon by the material properties and the cooling effectiveness of the first stage stator vanes and rotor blades, since these airfoils are exposed to the highest temperature flow.
Also in an industrial gas turbine engine, the life of a particular blade or vane is another important factor. When a turbine part is damaged from thermal or stress degradation, the engine cannot operate for a long period of time or the efficiency is decreased from a damaged part. The first stage turbine blade of an IGT engine is exposed to a stream of working fluid that is extremely hot (above 2000 degrees F.) and moving very quickly (above 500 ft/second). The rotor blades and stator vanes in this environment must tolerate not only extreme thermal loads, but also high-magnitude dynamic loads as well. As a result, these turbine parts are traditionally rigid, internally cooled structures that often include external thermal barrier coatings.
While robust architecture and thermal barrier coatings help the blades and vanes withstand external thermal and mechanical loads, they do not address all of the issues associated with exposure to the working fluid. For example, non-uniform temperature distribution between the cooled airfoil portions and relatively hot shroud portions introduces thermal gradients that produce internal thermal stresses. Cooling channel exits also produce localized thermal stresses, by inducing thermal gradients in the areas immediately surrounding the exits, as a result of sharp drops in temperature.
In these airfoils, thermal stress of an especially large magnitude occurs between the base portion and the platform of the rotor blade. The reason for this can be explained by the fact that since the moving blade has a smaller heat capacity than the platform, the temperature of the moving blade increases at a higher rate and within a shorter time period than that of the platform upon start of the gas turbine. On the other hand, the temperature of the moving blade falls at a higher rate and within a shorter time than that of the platform, whereby a large temperature difference occurs between the moving blade and the platform. This in turn generates thermal stress. Consequently, the base portion is shaped in the form of a curved surface conforming to the fillet ellipse to thereby reduce the thermal stress.
Recently, however, there is an increasing tendency to use a high temperature combustion gas to enhance the operating efficiency of the gas turbine. As a result, it becomes impossible to sufficiently suppress the thermal stress with only the base portion structure shaped in the form of the above mentioned fillet ellipse portion R, and cracks develop more frequently in the base portion where large thermal stress is induced. Under these circumstances, there is a demand for a structure of the blade base portion that is capable of reducing the thermal stress more effectively.
U.S. Pat. No. 6,481,967 B2 issued to Tomita et al on Nov. 19, 2002 and entitled GAS TURBINE MOVING BLADE shows a turbine rotor blade in FIGS. 1 and 2 with a row of trailing edge discharge cooling slots to provide cooling to the trailing edge and the fillet formed between the airfoil portion and the platform of the blade. High thermally induced stress is normally predicted at the junction of the blade trailing edge and the platform location. Also, due to the different effectiveness level of cooling mechanism used for the blade and platform and to the mass distribution between the blade and the platform, the thermally induced strain during transient cycle becomes much more severe. One method to alleviate this high thermal strain is by the use of a compounded fillet radii as shown in FIG. 3 which is disclosed in U.S. Pat. No. 6,851,924 B2 issued to Mazzola et al U.S. Pat. No. 6,851,924 B2 on Feb. 8, 2005 and entitled CRACK-RESISTANCE VANE SEGMENT MEMBER. As a result of this approach, the blade root section wall thickness must be increased which lowers the effectiveness of the trailing edge root section cooling slot. This results in a hotter trailing edge fillet metal temperature and a lower low cycle fatigue (LCF) capability.
It is an object of the present invention to provide for a turbine blade with a trailing edge root section cooling slot that will reduce the metal temperature in order to increase the LCF over the cited prior art references.